Airplanes have been most or totally built up from metallic components providing a good performance in terms of mechanical behaviour but, as a drawback, they also provide too much weight.
With the increase of competition among the airlines, airframe manufacturers search new ways of improving specific performances, meaning increasing or maintaining structural characteristics and decreasing weight against metallic builds.
One of the most important solutions is the use of composite fibre reinforced polymers (CFRP) for major structural parts, achieving important weight and operating costs savings. The first aircraft with a large CFRP composition is the Airbus 320, with more than 20%.
In summary, composites have been demonstrated to fulfil the following requirements:                Weight savings.        Be cost effective.        Meet structural requisite under aircraft conditions.        Beneficial cost/weight relation.        
The main structure for aircraft lifting surfaces consists of a leading edge, a torsion box, a trailing edge, a root joint and a tip. The torsion box consists of several structural elements: upper and lower skins stiffened by stringers on one hand, and spars and ribs on the other hand. Typically, the structural elements forming the torsion box are manufactured separately and are joined with the aid of complicated tooling to achieve the necessary tolerances, which are given by the aerodynamic, assembly and structural requirements. Nowadays, and particularly in the aeronautical industry, composite materials with an organic matrix and continuous fibres, especially CFRP (Carbon Fibre Reinforced Plastic), are widely used in a great variety of structural elements. For example, all the elements which make up a torsion box enumerated beforehand (ribs, stringers, spars and skins) can be manufactured using CFRP. Typically, of the different components which comprise a torsion box are manufactured separately and are thereafter assembled using rivets or other type of joining means.
There are several patents regarding integration of parts: U.S. Pat. No. 6,320,118B1 (Adhesively bonded joints in carbon fibre composite structures), U.S. Pat. No. 6,306,239B1 (Method of fabricating a stringer-stiffened shell structure using carbon reinforced composites), U.S. Pat. No. 4,749,155 (Method of making wing box cover panel), U.S. Pat. No. 5,454,895 (Process of manufacturing fibre reinforced structures suitable for aerodynamic applications), U.S. Pat. No. 5,817,269 (Composite fabricating method and tooling to improve part consolidation), all of them describing integration methods with a certain degree.
Another patent that describes a large degree of integration is the EP2153979A1, (Integrated multispar torsion box composite material), which also proposes a change in the typical arrangement of a torsion box formed by skin, stringers, front and rear spar and ribs. Nonetheless, this patent document concerns the integration of a multispar torsion box, with no ribs.
These solutions of the state of the art present the following technical problems:
Technical problems related to structures using rivets are, mainly:                the addition of weight to the resulting structures and        assembly time longer than desired.        
On the other hand, integrated structures also have problems such as:                They require very complex tooling, which sometimes even renders the process unprofitable.        it is desirable to save more time in the manufacture and the assembly of different parts of torsion boxes, especially those comprising ribs,        
There is a need for a solution providing a compromise between no integration and total integration so that the mentioned problems can be solved. It is desirable to have structures for which the number of rivets can be minimized by integrating different components which make up the structure in the least number of curing cycles possible. Even though there can be found, in the state of the art, solutions aiming to do so, these solutions still fail in providing the appropriate tooling in terms of complexity.
In this document the wording “composite material” is understood as any type of material, for example CFRP (Carbon Fibre Reinforced Polymers), which comprises two or more physically distinguishable parts and mechanically separable, the two or more parts not being able to dissolve among each other.
In the present description the following terms are defined as:                Co-curing: the process of joining two composite laminates provided in a fresh state by means of a single curing cycle. The resulting joint is certified for primary structures.        Co-bonding: the process of joining a composite laminate provided in a fresh state to a cured composite laminate by means of a curing cycle and the application of an adhesive along the joining surface of the laminates. The resulting joint is certified for primary structures.        Bonding: the process of joining two cured composite parts by means of an adhesive material. It is not certified for joining primary structures.        Mechanical bonding the process of joining two parts by means of fastening means, such as rivets or bolts The resulting joint is certified for primary structures.        